Turbine engine shroud assembly

ABSTRACT

Disclosed herein is an interlocking shroud assembly for a turbine engine with a plurality of radially extending, circumferentially spaced airfoils terminating in a shroud element and having opposing radial sides with first and second interlock elements. Further provided is a method of forming a shroud about a plurality of rotating blades in a turbine engine.

BACKGROUND

Turbine engines, and particularly gas or combustion turbine engines, arerotary engines that extract energy from a flow of pressurized combustedgases passing through the engine onto a multitude of rotating turbineblades.

The rotating turbine blades can be supported by shrouds that areinterlocked to form a circumferential casing to the turbine. A Z-shapedinterlock is a typical configuration choice for a shrouded bladeassembly which requires a pre-twist during manufacturing and assembly.Eliminating the pre-twist while maintaining an interlock configurationwould be beneficial for shroud assembly manufacturing.

BRIEF DESCRIPTION

In one aspect, embodiments of relate to a turbine engine comprising arotor having a plurality of radially extending airfoils spacedcircumferentially about the rotor, with the airfoils terminating in atip, and a shroud assembly circumscribing the airfoils and comprising ashroud element mounted to each tip and having opposing radial sides withfirst and second interlock elements, wherein the first interlock elementof one shroud element mates with a second interlock element of acircumferentially adjacent element to form a plurality of interlocksbetween adjacent shroud elements about the circumference of theairfoils.

In another aspect, embodiments relate to an interlocking shroud assemblyfor a turbine engine comprising a plurality of radially extending,circumferentially spaced airfoils terminating in a shroud element andhaving opposing radial sides with first and second interlock elements,which interlock to form interlocks between circumferentially adjacentairfoils.

In yet another aspect, embodiments relate to a method of forming ashroud about a plurality of rotating blades in a turbine enginescomprising forming an interlock between circumferentially adjacent tipsof the blades and preloading the interlock.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawings:

FIG. 1 is a schematic cross-sectional diagram of a turbine engine for anaircraft.

FIG. 2 is an assembled plurality of airfoils.

FIG. 3 is a perspective view of a shroud element.

FIG. 4 is another perspective view of a shroud element.

FIG. 5 is an illustration of a shroud assembly.

FIG. 6 is a cross-sectional view of the shroud assembly of FIG. 5.

FIG. 7 is a cross-sectional view of a second embodiment of the shroudassembly of FIG. 5.

FIG. 8 is a single airfoil assembly.

DETAILED DESCRIPTION

The described embodiments of the present invention are directed to ashroud assembly for an airfoil. For purposes of illustration, thepresent invention will be described with respect to the turbine for anaircraft turbine engine. It will be understood, however, that theinvention is not so limited and may have general applicability within anengine, including compressors, as well as in non-aircraft applications,such as other mobile applications and non-mobile industrial, commercial,and residential applications.

As used herein, the term “forward” or “upstream” refers to moving in adirection toward the engine inlet, or a component being relativelycloser to the engine inlet as compared to another component. The term“aft” or “downstream” used in conjunction with “forward” or “upstream”refers to a direction toward the rear or outlet of the engine relativeto the engine centerline.

Additionally, as used herein, the terms “radial” or “radially” refer toa dimension extending between a center longitudinal axis of the engineand an outer engine circumference.

All directional references (e.g., radial, axial, proximal, distal,upper, lower, upward, downward, left, right, lateral, front, back, top,bottom, above, below, vertical, horizontal, clockwise, counterclockwise,upstream, downstream, aft, etc.) are only used for identificationpurposes to aid the reader's understanding of the present invention, anddo not create limitations, particularly as to the position, orientation,or use of the invention. Connection references (e.g., attached, coupled,connected, and joined) are to be construed broadly and can includeintermediate members between a collection of elements and relativemovement between elements unless otherwise indicated. As such,connection references do not necessarily infer that two elements aredirectly connected and in fixed relation to one another. The exemplarydrawings are for purposes of illustration only and the dimensions,positions, order and relative sizes reflected in the drawings attachedhereto can vary.

FIG. 1 is a schematic cross-sectional diagram of a turbine engine 10 foran aircraft. The engine 10 has a generally longitudinally extending axisor centerline 12 extending forward 14 to aft 16. The engine 10 includes,in downstream serial flow relationship, a fan section 18 including a fan20, a compressor section 22 including a booster or low pressure (LP)compressor 24 and a high pressure (HP) compressor 26, a combustionsection 28 including a combustor 30, a turbine section 32 including a HPturbine 34, and a LP turbine 36, and an exhaust section 38.

The fan section 18 includes a fan casing 40 surrounding the fan 20. Thefan 20 includes a plurality of fan blades 42 disposed radially about thecenterline 12. The HP compressor 26, the combustor 30, and the HPturbine 34 form a core 44 of the engine 10, which generates combustiongases. The core 44 is surrounded by core casing 46, which can be coupledwith the fan casing 40.

A HP shaft or spool 48 disposed coaxially about the centerline 12 of theengine 10 drivingly connects the HP turbine 34 to the HP compressor 26.A LP shaft or spool 50, which is disposed coaxially about the centerline12 of the engine 10 within the larger diameter annular HP spool 48,drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20.

The LP compressor 24 and the HP compressor 26 respectively include aplurality of compressor stages 52, 54, in which a set of compressorblades 56, 58 rotate relative to a corresponding set of staticcompressor vanes 60, 62 (also called a nozzle) to compress or pressurizethe stream of fluid passing through the stage. In a single compressorstage 52, 54, multiple compressor blades 56, 58 can be provided in aring and can extend radially outwardly relative to the centerline 12,from a blade platform to a blade tip, while the corresponding staticcompressor vanes 60, 62 are positioned upstream of and adjacent to therotating blades 56, 58. It is noted that the number of blades, vanes,and compressor stages shown in FIG. 1 were selected for illustrativepurposes only, and that other numbers are possible.

The blades 56, 58 for a stage of the compressor can be mounted to a disk59, which is mounted to the corresponding one of the HP and LP spools48, 50, with each stage having its own disk 59, 61. The vanes 60, 62 fora stage of the compressor can be mounted to the core casing 46 in acircumferential arrangement.

The HP turbine 34 and the LP turbine 36 respectively include a pluralityof turbine stages 64, 66, in which a set of turbine blades 68, 70 arerotated relative to a corresponding set of static turbine vanes 72, 74(also called a nozzle) to extract energy from the stream of fluidpassing through the stage. In a single turbine stage 64, 66, multipleturbine vanes 72, 74 can be provided in a ring and can extend radiallyoutwardly relative to the centerline 12, while the correspondingrotating blades 68, 70 are positioned downstream of and adjacent to thestatic turbine vanes 72, 74 and can also extend radially outwardlyrelative to the centerline 12, from a blade platform to a blade tip. Itis noted that the number of blades, vanes, and turbine stages shown inFIG. 1 were selected for illustrative purposes only, and that othernumbers are possible.

The blades 68, 70 for a stage of the turbine can be mounted to a disk71, which is mounted to the corresponding one of the HP and LP spools48, 50, with each stage having its own disk 71, 73. The vanes 72, 74 fora stage of the compressor can be mounted to the core casing 46 in acircumferential arrangement.

The portions of the engine 10 mounted to and rotating with either orboth of the spools 48, 50 are also referred to individually orcollectively as a rotor 53. The stationary portions of the engine 10including portions mounted to the core casing 46 are also referred toindividually or collectively as a stator 63.

In operation, the airflow exiting the fan section 18 is split such thata portion of the airflow is channeled into the LP compressor 24, whichthen supplies pressurized ambient air 76 to the HP compressor 26, whichfurther pressurizes the ambient air. The pressurized air 76 from the HPcompressor 26 is mixed with fuel in the combustor 30 and ignited,thereby generating combustion gases. Some work is extracted from thesegases by the HP turbine 34, which drives the HP compressor 26. Thecombustion gases are discharged into the LP turbine 36, which extractsadditional work to drive the LP compressor 24, and the exhaust gas isultimately discharged from the engine 10 via the exhaust section 38. Thedriving of the LP turbine 36 drives the LP spool 50 to rotate the fan 20and the LP compressor 24.

A remaining portion of the airflow 78 bypasses the LP compressor 24 andengine core 44 and exits the engine assembly 10 through a stationaryvane row, and more particularly an outlet guide vane assembly 80,comprising a plurality of airfoil guide vanes 82, at the fan exhaustside 84. More specifically, a circumferential row of radially extendingairfoil guide vanes 82 are utilized adjacent the fan section 18 to exertsome directional control of the airflow 78.

Some of the ambient air supplied by the fan 20 can bypass the enginecore 44 and be used for cooling of portions, especially hot portions, ofthe engine 10, and/or used to cool or power other aspects of theaircraft. In the context of a turbine engine, the hot portions of theengine are normally the combustor 30 and components downstream of thecombustor 30, especially the turbine section 32, with the HP turbine 34being the hottest portion as it is directly downstream of the combustionsection 28. Other sources of cooling fluid can be, but is not limitedto, fluid discharged from the LP compressor 24 or the HP compressor 26.This fluid can be bleed air 77 which can include air drawn from the LPor HP compressors 24, 26 that bypasses the combustor 30 as coolingsources for the turbine section 32. This is a common engineconfiguration, not meant to be limiting.

FIG. 2 illustrates a plurality of radially extending circumferentiallyspaced airfoils, or blades 70, with each blade 70 extending from a root88 and terminating in a tip (FIG. 3) arranged in a circumferential rowand supported by an arcuate inner band 96 and an arcuate outer band 98.The arcuate outer band 98 comprises a shroud assembly 100 made up ofseparate individual shroud elements 102, having opposing radial sides107, 109, which together circumscribe the blades 70.

Each shroud element 102 as depicted in FIGS. 3 and 4 is integrallyformed with the blade 70 at the tip 90 and comprises a flange 104. Theflange 104 includes a first and a second radial edge 106, 108 and a seat110 that projects circumferentially beyond the first radial edge 106illustrated in FIG. 3. The seat 110 is formed by the first radial edge106 and the flange 104 in a portion 112 of the flange 104 that iscircumferentially outboard of the first radial edge 106 illustrated inFIG. 4.

FIG. 5 illustrates an airfoil assembly 114 in which the shroud element102 is integrally formed with the blade 70 wherein the blade terminatesin a dovetail 116. The dovetail 116 is formed to mount to the rotor 53.When assembled, the blade 70 is sprung 118 to apply a preload to theinterlocks. The blade 70 can be sprung as shown by solid line 118 from apredominantly parallel position 120 with respect to a neutral axis 122of the airfoil assembly 114 to a bowed position 124 when interlocked.The blade 70 can also be sprung 118 from an initial position ofpredominantly bowed 126 to a parallel position 128 when interlocked.

Regardless of the initial or final positions of the blade 70, the finalposition 124, 128 will cause the second radial edge 108 bias outwardlyand the seat to bias inwardly. This bias is caused by the compressiveforce F_(C) from the dovetail 116 which translates to an upward force F₂at the second radial edge 108 and a downward force F₁ from the seat 110.

Circumferentially adjacent shroud elements 102 interlock togetherforming a plurality of interlocks 130 between adjacent shroud elements102 to form the shroud assembly 100 as illustrated in FIG. 6. Across-sectional view of an exemplary embodiment of the shroud assembly100 of FIG. 6 is shown in FIG. 7 where a first flange 132 having a firstinterlock element 134 mates with a second flange 136 having a secondinterlock element 138 where the first flange 132 overlies and abuts thesecond flange 136.

When the shroud assembly 100 is assembled, the second radial edge 108 ofthe second flange 136 will bias toward the first flange 132 due to theforces F₁ and F₂. This bias enables friction forces to form between thefirst and second interlock element surfaces that bond each shroudelement 102 to the next radially adjacent shroud element 102.

A second embodiment of the shroud assembly is contemplated in FIG. 8.The second embodiment is similar to the first embodiment, therefore,like parts will be identified with like numerals increasing by 100respectively, with it being understood that the description of the likeparts of the first embodiment applies to the additional embodiments,unless otherwise noted.

In the second embodiment a first interlock element 234 formed on a firstflange 232 includes an angled seat 210 formed to receive an angledsecond interlock element 238 formed on a second flange 236. The flange204 of each shroud element 202 therefore includes an angled seat 210 andan angled portion 240 formed to fit into the angled seat 210. Whileillustrated as two ramps 242, 244 forming an apex 246, the angled seat210 and angled portion 240 can be any shape where the first interlockelement 234 is formed to receive the second interlock element 238.

A method of forming a shroud assembly, comprising a shroud elementintegrally formed with a blade, about a plurality of rotating blades ina turbine engines includes forming an interlock betweencircumferentially adjacent tips of the blades and preloading theinterlock. The preloading of the interlock where an interlock element ismade to bias towards another interlock element.

The embodiments described herein have benefits regarding production,performance, and damping capability. Prior art for shroud bladeassemblies include Z-shaped interlocks. Implementing airfoil and shroudradial bending ensures contact between interlock elements to achieveouter band preload at operating conditions without typical torsionalbending used in Z-shaped shroud design which requires a pre-twist. Thistype of bending also only requires blade balancing for centrifugalforces and improves damping at blade resonant vibrations due toincreased contact areas between the interlock element surfaces. Thisincrease in contact area also provides for a reduction in outer flowpathleakages. The simplified shape of design and removing a need for apre-twist, eases manufacturing.

It should be appreciated that application of the disclosed design is notlimited to turbine engines with fan and booster sections, but isapplicable to turbojets and turbo engines as well.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they have structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A turbine engine comprising: a rotor having aplurality of radially extending airfoils spaced circumferentially aboutthe rotor, with the airfoils terminating in a tip; and a shroud assemblycircumscribing the airfoils and comprising a shroud element mounted toeach tip and having opposing radial sides with first and secondinterlock elements; wherein the first interlock element of one shroudelement mates with a second interlock element of a circumferentiallyadjacent element to form a plurality of interlocks between adjacentshroud elements about the circumference of the airfoils.
 2. The turbineengine of claim 1, wherein the shroud element is integrally formed withthe airfoil.
 3. The turbine engine of claim 2, wherein the airfoil is ablade.
 4. The turbine engine of claim 3, wherein the blade terminates ina dove tail opposite the tip and the dove tail is mounted to the rotor.5. The turbine engine of claim 1, wherein the airfoil is sized such thatthe airfoil is sprung when interlocked with adjacent airfoils to apply apreload to the interlocks.
 6. The turbine engine of claim 1, wherein thefirst interlock comprises a first flange, the second interlock comprisesa second flange, and the first flange overlies and abuts the secondflange.
 7. The turbine engine of claim 6, wherein the first interlockfurther comprises a seat spaced circumferentially from the first flangeand the second flange sits within the seat.
 8. The turbine engine ofclaim 7, wherein the seat is formed by a first radial edge and the firstflange.
 9. The turbine engine of claim 8, wherein the first flangeprojects circumferentially beyond the first radial edge.
 10. The turbineengine of claim 9, wherein the first flange is circumferentiallyoutboard of the first radial edge.
 11. The turbine engine of claim 10,wherein the second flange is a second radial edge of an adjacent shroudelement.
 12. The turbine engine of claim 11, wherein the airfoil issized such that the airfoil is sprung when interlocked with adjacentairfoils to apply a preload to the interlocks causing the second radialedge to bias toward the first flange.
 13. An interlocking shroudassembly for a turbine engine comprising a plurality of radiallyextending, circumferentially spaced airfoils terminating in a shroudelement and having opposing radial sides with first and second interlockelements, which interlock to form interlocks between circumferentiallyadjacent airfoils.
 14. The interlocking shroud assembly of claim 13,wherein the airfoil is sized such that the airfoil is sprung wheninterlocked with adjacent airfoils to apply a preload to the interlocks.15. The interlocking shroud assembly of claim 14, wherein the firstinterlock comprises a first flange, the second interlock comprises asecond flange and the first flange overlies and abuts the second flange.16. The interlocking shroud assembly of claim 15, wherein the firstinterlock further comprises a seat spaced circumferentially from thefirst flange and the second flange sits within the seat.
 17. Theinterlocking shroud assembly of claim 16, wherein the seat is formed bya first radial edge and the first flange.
 18. The interlocking shroudassembly of claim 17, wherein the first flange projectscircumferentially beyond the first radial edge.
 19. The interlockingshroud assembly of claim 18, wherein the first flange iscircumferentially outboard of the first radial edge.
 20. Theinterlocking shroud assembly of claim 19, wherein the second flange is asecond radial edge of an adjacent shroud element.
 21. A method offorming a shroud about a plurality of rotating blades in a turbineengines comprising: forming an interlock between circumferentiallyadjacent tips of the blades; and preloading the interlock.
 22. Themethod of claim 21, wherein forming the interlock comprises forming aninterlock on circumferentially opposite sides of the tip of the blade.23. The method of claim 21, wherein preloading the interlock comprisesspringing the blade.